The present invention relates to turbine engines, and more particularly to an inflatable bleed valve for a low pressure compressor for a turbine engine, such as a tip turbine engine.
An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan, a low pressure compressor, a middle core engine, and an aft low pressure turbine, all located along a common longitudinal axis. A high pressure compressor and a high pressure turbine of the core engine are interconnected by a high pressure shaft. The high pressure compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in a combustor, where it is ignited to form a high energy gas stream. The gas stream flows axially aft to rotatably drive the high pressure turbine, which rotatably drives the high pressure compressor via the high pressure shaft. The gas stream leaving the high pressure turbine is expanded through the low pressure turbine, which rotatably drives the bypass fan and low pressure compressor via a low pressure shaft.
Although highly efficient, conventional turbofan engines operate in an axial flow relationship. The axial flow relationship results in a relatively complicated elongated engine structure of considerable length relative to the engine diameter. This elongated shape may complicate or prevent packaging of the engine into particular applications.
A recent development in gas turbine engines is the tip turbine engine. Tip turbine engines may include a low pressure axial compressor directing core airflow into hollow fan blades. The hollow fan blades operate as a centrifugal compressor when rotating. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor, where it is ignited to form a high energy gas stream which drives the turbine that is integrated onto the tips of the hollow bypass fan blades for rotation therewith as generally disclosed in U.S. Patent Application Publication Nos.: 20030192303; 20030192304; and 20040025490. The tip turbine engine provides a thrust-to-weight ratio equivalent to or greater than conventional turbofan engines of the same class, but within a package of significantly shorter length.
The compressors for turbine engines are designed at the maximum power point. When operating at partial power points it sometimes becomes necessary to bleed air form the back of the compressor for stage matching reasons. At times, the rear compressor stages cannot handle the amount of flow that the front stages are pumping. To match flow, some air is bled off to reduce the flow entering the rear stages. Turbine engines may also use bleed air internally for accessory functions. Some bleed air may be discharged radially out through some of the turbine blades or stators for cooling purposes.
The compressor of a conventional turbine engine includes a bleed valve assembly including a rotating and translating ring with linkages. A large hydraulic actuator is disposed immediately proximate the bleed valve for selectively opening and closing the bleed valve. These bleed valve assemblies are large, heavy and complex. Moreover, these bleed valve assemblies are not easily packaged into the low pressure axial compressors for tip turbine engines. Conventional bleed valves like this are also radially inward of the bypass flow; however, the low compressor in conventional engines dips radially inward at the aft end of providing the room needed for the bleed valve. This is not true on the tip turbine engine.